Gas turbine engine airfoil

ABSTRACT

An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil geometry corresponds to axial leading and trailing edge curves and an axial stacking offset curve. The airfoil extends from a root and a zero axial reference point corresponds to axial center of the root. X LE  corresponds to an axial distance from a leading edge to the reference point at a given span position. X TE  corresponds to a axial distance from a trailing edge to the reference point at a given span position. X d  corresponds to an axial stacking offset at a given span position. (X LE −X d )/(X d −X TE ) at 100% span position is about 1 and (X LE −X d )/(X d −X TE ) at 90% span position is about 1.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Application No.61/941,768, which was filed on Feb. 19, 2014 and is incorporated hereinby reference.

BACKGROUND

This disclosure relates generally to an airfoil for gas turbine engines,and more particularly to axial projection relative to span.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. Air entering the compressor section is compressed and deliveredinto the combustor section where it is mixed with fuel and ignited togenerate a high-speed exhaust gas flow. The high-speed exhaust gas flowexpands through the turbine section to drive the compressor and the fansection. The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

The propulsive efficiency of a gas turbine engine depends on manydifferent factors, such as the design of the engine and the resultingperformance debits on the fan that propels the engine. As an example,the fan may rotate at a high rate of speed such that air passes over thefan airfoils at transonic or supersonic speeds. The fast-moving aircreates flow discontinuities or shocks that result in irreversiblepropulsive losses. Additionally, physical interaction between the fanand the air causes downstream turbulence and further losses. Althoughsome basic principles behind such losses are understood, identifying andchanging appropriate design factors to reduce such losses for a givenengine architecture has proven to be a complex and elusive task.

SUMMARY

In one exemplary embodiment, an airfoil for a turbine engine includespressure and suction sides that extend in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip. The airfoil geometry corresponds to axial leading andtrailing edge curves and an axial stacking offset curve. The airfoilextends from a root and a zero axial reference point corresponds toaxial center of the root. X_(LE) corresponds to an axial distance from aleading edge to the reference point at a given span position. X_(TE)corresponds to a axial distance from a trailing edge to the referencepoint at a given span position. X_(d) corresponds to an axial stackingoffset at a given span position. (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100%span position is about 1 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% spanposition is about 1.

In a further embodiment of the above, (X_(LE)−X_(d))/(X_(d)−X_(TE)) at60% span position is about 1.2 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50%span position about 0.92.

In a further embodiment of any of the above,(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position is about 0.92 and(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position is about 0.92.

In a further embodiment of any of the above,(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% span position is about 0.92 and(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% span position is about 0.92.

In a further embodiment of any of the above,(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% span position is about 0.89 and(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% span position is about 1.

In a further embodiment of any of the above,(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position about 1 and(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position is about 1.

In a further embodiment of any of the above, the airfoil is a fan bladefor a gas turbine engine.

In a further embodiment of any of the above,(X_(LE)−X_(d))/(X_(d)−X_(TE)) has a tolerance of +/−0.05.

In another exemplary embodiment, a gas turbine engine includes acombustor section arranged between a compressor section and a turbinesection, a fan section that has an array of twenty-six or fewer fanblades that have a low fan pressure ratio of less than 1.55 and a gearedarchitecture coupling the fan section to the turbine section or thecompressor section. The fan blades include an airfoil that has pressureand suction sides. The airfoil extends in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip. The airfoil geometry corresponds to axial leading andtrailing edge curves and an axial stacking offset curve. The airfoilextends from a root and a zero axial reference point corresponds to anaxial center of the root. X_(LE) corresponds to an axial distance from aleading edge to the reference point at a given span position. X_(TE)corresponds to an axial distance from a trailing edge to the referencepoint at a given span position. X_(d) corresponds to an axial stackingoffset at a given span position. (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100%span position is about 1 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% spanposition is about 1.

In a further embodiment of any of the above,(X_(LE)−X_(d))/(X_(d)−X_(TE)) has a tolerance of +/−0.05.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment.

FIG. 2A is a perspective view of a portion of a fan section.

FIG. 2B is a schematic cross-sectional view of the fan section.

FIG. 2C is a cross-sectional view a fan blade taken along line 2C-2C inFIG. 2B.

FIG. 3A is a schematic view of fan blade span positions.

FIG. 3B is a schematic view of a cross-section of a fan blade section ata particular span position and its axial twist and chord parameters.

FIG. 4A illustrates a relationship between axial leading edge position,axial stacking offset position and axial trailing edge position relativeto span position for a set of first example airfoils.

FIG. 4B illustrates a relationship between axial leading edge position,axial stacking offset position and axial trailing edge position relativeto span position for a set of second example airfoils.

FIG. 4C illustrates a relationship between axial leading edge position,axial stacking offset position and axial trailing edge position relativeto span position for a set of third example airfoils.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures. That is, the disclosedairfoils may be used for engine configurations such as, for example,direct fan drives, or two- or three-spool engines with a speed changemechanism coupling the fan with a compressor or a turbine sections.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisX which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption ('TSFC')”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.55. Inanother non-limiting embodiment the low fan pressure ratio is less thanabout 1.45. In another non-limiting embodiment the low fan pressureratio is from 1.1 to 1.45. “Low corrected fan tip speed” is the actualfan tip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram° R)/(518.7° R)]^(0.5). The “Low corrected fan tipspeed” as disclosed herein according to one non-limiting embodiment isless than about 1150 ft/second. The “low corrected fan tip speed” asdisclosed herein according to another non-limiting embodiment is lessthan about 1200 ft/second.

Referring to FIG. 2A-2C, the fan blade 42 is supported by a fan hub 60that is rotatable about the axis X. Each fan blade 42 includes anairfoil 64 extending in a radial span direction R from a root 62 to atip 66. A 0% span position corresponds to a section of the airfoil 64 atthe inner flow path (e.g., a platform), and a 100% span positioncorresponds to a section of the airfoil 64 at the tip 66.

The root 62 is received in a correspondingly shaped slot in the fan hub60. The airfoil 64 extends radially outward of the platform, whichprovides the inner flow path. The platform may be integral with the fanblade or separately secured to the fan hub, for example. A spinner 66 issupported relative to the fan hub 60 to provide an aerodynamic innerflow path into the fan section 22.

The airfoil 64 has an exterior surface 76 providing a contour thatextends from a leading edge 68 aftward in a chord-wise direction H to atrailing edge 70, as shown in FIG. 2C. Pressure and suction sides 72, 74join one another at the leading and trailing edges 68, 70 and are spacedapart from one another in an airfoil thickness direction T. An array ofthe fan blades 42 are positioned about the axis X in a circumferentialor tangential direction Y. Any suitable number of fan blades may be usedin a given application.

The exterior surface 76 of the airfoil 64 generates lift based upon itsgeometry and directs flow along the core flow path C. The fan blade 42may be constructed from a composite material, or an aluminum alloy ortitanium alloy, or a combination of one or more of these.Abrasion-resistant coatings or other protective coatings may be appliedto the fan blade 42.

One characteristic of fan blade performance relates to the fan blade'saxial stacking offset and leading and trailing edge positions (Xdirection) relative to a particular span position (R direction).Referring to FIG. 3A, span positions a schematically illustrated from 0%to 100% in 10% increments. Each section at a given span position isprovided by a conical cut that corresponds to the shape of the core flowpath, as shown by the large dashed lines. In the case of a fan bladewith an integral platform, the 0% span position corresponds to theradially innermost location where the airfoil meets the fillet joiningthe airfoil to the platform. In the case of a fan blade without anintegral platform, the 0% span position corresponds to the radiallyinnermost location where the discrete platform meets the exteriorsurface of the airfoil. In addition to varying with span, axialprojection varies between a hot, running condition and a cold, static(“on the bench”) condition.

The X_(CG) corresponds to the location of the center of gravity for aparticular section at a given span location relative to a referencepoint 80 in the X direction, as shown in FIG. 3B. The center of gravityassumes a homogenous material. The reference point 80 is the axialcenter of the root, and X_(d) corresponds to the axial distance from thereference point 80 to the center of gravity.

A positive X value corresponds to the aftward direction along theengine's axis of rotation. A negative X value corresponds to the forwarddirection along the engine's axis of rotation.

The axial leading edge location is arranged at the leading edge 68 for aparticular section at a given span location relative to the referencepoint 80 in the X direction, as shown in FIG. 3B. X_(LE) corresponds tothe axial distance from the reference point 80 to the axial leading edgelocation at a given span location.

The axial trailing edge location is arranged at the trailing edge 70 fora particular section at a given span location relative to the referencepoint 80 in the X direction. X_(TE) corresponds to the axial distancefrom the reference point 80 to the axial trailing edge location at agiven span location.

The axial changes in fan blade axial projection at various spanpositions can be expressed using the differences X_(LE)−X_(d) andX_(d)−X_(TE), which are tangential distances between the locations.These differences can be used to provide non-dimensional ratiosindicative of desired airfoil characteristics.

In one prior art airfoil, (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100% spanposition is about 0.89 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% spanposition is about 0.91; (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% spanposition is about 1.07 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% spanposition about 0.88; and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% spanposition is about 0.94 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% spanposition is about 0.88.

Example relationships between the axial projection (X) relative to thespan position (AVERAGE SPAN %) are shown in FIGS. 4A-4C for severalexample fan blades, each represented by a curve. Only one curve in eachgraph is discussed for simplicity. The airfoil geometry corresponds toaxial leading and trailing edge curves and an axial stacking offsetcurve, and a greatest differential between the axial leading edgeposition and the axial trailing edge position along the curves is in therange of 30-50% span position. “About” used in relation to the(X_(LE)−X_(d))/(X_(d)−X_(TE)) ratios means +/−0.10 in one example, and+/−0.05 in another example.

Referring to FIG. 4A, (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100% spanposition is about 1 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% spanposition is about 1; (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% span positionis about 1.2 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% span positionabout 0.92; and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position isabout 0.92 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position isabout 0.92.

Referring to FIG. 4B, (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100% spanposition is about 1 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% spanposition is about 1; (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% span positionis about 0.92 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% span position isabout 0.92; and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position isabout 0.92 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position isabout 0.92.

Referring to FIG. 4C, (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100% spanposition is about 1 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% spanposition is about 1; (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% span positionis about 0.89 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% span position isabout 1; and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position about 1and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position is about 1.

The axial leading and trailing edge positions and axial stacking offsetsin a hot, running condition along the span of the airfoils 64 relate tothe contour of the airfoil and provide necessary fan operation in cruiseat the lower, preferential speeds enabled by the geared architecture 48in order to enhance aerodynamic functionality and thermal efficiency. Asused herein, the hot, running condition is the condition during cruiseof the gas turbine engine 20. For example, the axial leading andtrailing edge positions and axial stacking offsets in the hot, runningcondition can be determined in a known manner using numerical analysis,such as finite element analysis

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil for a turbine engine comprising:pressure and suction sides extending in a radial direction from a 0%span position at an inner flow path location to a 100% span position atan airfoil tip, wherein the airfoil geometry corresponds to axialleading and trailing edge curves and an axial stacking offset curve,wherein the airfoil extends from a root, and a zero axial referencepoint corresponds to axial center of the root, X_(LE) corresponds to aaxial distance from a leading edge to the reference point at a givenspan position, X_(TE) corresponds to a axial distance from a trailingedge to the reference point at a given span position, X_(d) correspondsto an axial stacking offset at a given span position, wherein(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 100% span position is 1+/−0.10 and(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 90% span position is 1+/−0.10.
 2. Theairfoil according to claim 1, wherein (X_(LE)−X_(d))/(X_(d)−X_(TE)) at60% span position is 1.2+/−0.10 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50%span position is 0.92+/−0.10.
 3. The airfoil according to claim 1,wherein (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position is0.92+/−0.10 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position is0.92+/−0.10.
 4. The airfoil according to claim 1, wherein(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 60% span position is 0.92+/−0.10 and(X_(LE)−X_(d))/(X_(d)−X_(TE)) at 50% span position is 0.92+/−0.10. 5.The airfoil according to claim 1, wherein (X_(LE)−X_(d))/(X_(d)−X_(TE))at 60% span position is 0.89+/−0.10 and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at50% span position is 1+/−0.10.
 6. The airfoil according to claim 1,wherein (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 40% span position is 1+/−0.10and (X_(LE)−X_(d))/(X_(d)−X_(TE)) at 20% span position is 1+/−0.10. 7.The airfoil according to claim 1, wherein the airfoil is a fan blade fora gas turbine engine.
 8. The airfoil according to claim 1, wherein(X_(LE)−X_(d))/(X_(d)−X_(TE)) has a tolerance of +/−0.05.
 9. A gasturbine engine comprising: a combustor section arranged between acompressor section and a turbine section; a fan section having an arrayof twenty-six or fewer fan blades, wherein the fan section has a low fanpressure ratio of less than 1.55; a geared architecture coupling the fansection to the turbine section or the compressor section; and whereinthe fan blades include an airfoil having pressure and suction sides, theairfoil extends in a radial direction from a 0% span position at aninner flow path location to a 100% span position at an airfoil tip,wherein the airfoil geometry corresponds to axial leading and trailingedge curves and an axial stacking offset curve, wherein the airfoilextends from a root, and a zero axial reference point corresponds toaxial center of the root, X_(LE) corresponds to an axial distance from aleading edge to the reference point at a given span position, X_(TE)corresponds to an axial distance from a trailing edge to the referencepoint at a given span position, X_(d) corresponds to an axial stackingoffset at a given span position, wherein (X_(LE)−X_(d))/(X_(d)−X_(TE))at 100% span position is 1+/−0.10 and (X_(LE) X_(d))/(X_(d)−X_(TE)) at90% span position is 1+/−0.10.
 10. The gas turbine engine according toclaim 9, wherein (X_(LE)−X_(d))/(X_(d)−X_(TE)) has a tolerance of+/−0.05.